The present invention relates to components designed to operate at high temperatures. More particularly, this invention relates to methods for repair and manufacture of blades for gas turbine engines, and the articles made and repaired from the use of these methods.
In a gas turbine engine, compressed air is mixed with fuel in a combustor and ignited, generating a flow of hot combustion gases through one or more turbine stages that extract energy from the gas, producing output power. Each turbine stage includes a stator nozzle having vanes which direct the combustion gases against a corresponding row of turbine blades extending radially outwardly from a blade root, where a dovetail joint attaches the blade to a supporting rotor disk, to a blade tip at the opposite end. The blades are subject to substantial heat load, and, because the efficiency of a gas turbine engine is proportional to gas temperature, the continuous demand for efficiency improvements translates to a demand for blades that are capable of withstanding higher temperatures for longer service times.
Gas turbine blades are usually made of superalloys and are often cooled by means of internal cooling chambers and the addition of coatings, including thermal barrier coatings (TBC's) and environmentally resistant coatings, to their external surfaces. The term “superalloy” is usually intended to embrace iron-, cobalt-, or nickel-based alloys, which include one or more other elements including such non-limiting examples as aluminum, tungsten, molybdenum, titanium, and iron. The internal air cooling of turbine blades is often accomplished via a complex cooling scheme in which cooling air flows through channels within the blade (“internal cooling channels”) and is then discharged through a configuration of cooling holes at the blade surface. Convection cooling occurs within the blade from heat transfer to the cooling air as it flows through the internal air cooling channels. In more complex configurations, fine internal orifices are often provided to direct cooling air flow directly against inner surfaces of the blade to achieve what is referred to as impingement cooling, while film cooling is often accomplished at the blade surface by configuring the cooling holes to discharge the cooling air flow across the blade surface so that the surface is protected from direct contact with the surrounding hot gases within the engine. TBC's comprise at least a layer of thermally insulating ceramic and often include one or more layers of metal-based, oxidation-resistant materials (“environmentally resistant coatings”) underlying the insulating ceramic for enhanced protection of the blade. Environmentally resistant coatings are also frequently used without a TBC topcoat. Technologies such as coatings and internal cooling have effectively enhanced the performance of turbine blades, but material degradation problems persist in turbine blades due to locally aggressive conditions in areas such as blade tips.
A considerable amount of cooling air is often required to sufficiently lower the surface temperature of a blade. However, the casting process and the cores required to form the cooling channels limit the complexity of the cooling scheme that can be formed within a blade at the blade tip. The resulting restrictions in cooling airflow often promote higher local temperatures in this area relative to those existing in other locations on a given blade. In typical jet engines, for example, bulk average blade temperatures range between about 900° C. to about 1000° C., while blade tip surfaces often reach bout 1100° C. or more. Maximum surface temperatures are expected in future applications to be over about 1300° C. Of particular concern is the combination of stress with temperature, because metals, including alloys used to make gas turbine blades, tend to become weaker, or more easily deformed, as temperatures increase. Thus, while stress of a certain level operating on a cooler section of a blade may have little effect on performance, the same stress level may be beyond the performance capability of the material at hotter locations as described above. At such elevated temperatures, materials are more susceptible to damage due to a number of phenomena, including diffusion-controlled deformation (“creep”), cyclic loading and unloading (“fatigue”), chemical attack by the hot gas flow (“oxidation”), wear from rubbing contact between blade tips and turbine shrouds, wear from the impact of particles entrained in the gas flow (“erosion”), and others.
Damage to blades, particularly at blade tips, leads to degradation of turbine efficiency. As blade tips are deformed, oxidized, or worn away, gaps between the blade tip and the turbine shroud become excessively wide, allowing gas to leak through the turbine stages without the flow of the gas being converted into mechanical energy. When efficiency drops below specified levels, the turbine must be removed from service for overhaul and refurbishment. A significant portion of this refurbishment process is directed at the repair of blade tips.
In current practice, the original blade tip material is made of the same material as the rest of the original blade, often a superalloy based on nickel or cobalt. Because this material was selected to balance the design requirements of the entire blade, it is generally not optimized to meet the special local requirements demanded by conditions at the blade tip. The performance of alloys commonly used for repair is comparable or inferior to that of the material of the original component, depending upon the microstructure, defect density, and chemistry of the repair material. For example, many turbine blades are made using alloys that have been directionally solidified. The directional solidification process manipulates the orientation of metal crystals, or grains, as the alloy is solidified from the molten state, aligning the grains in one selected primary direction. The resultant alloy has enhanced resistance to creep and fatigue during service when compared to conventionally processed materials. Advanced applications employ alloys made of a single crystal for even further improvements in high-temperature creep and fatigue behavior. However, when blade tips are repaired by some conventional processes, using build-up of weld filler material, the resulting microstructure of the repair is typical of welded material, not directionally solidified or single-crystalline. Other repair methods, such as applying powder mixtures wherein one powder melts and densifies the repaired area during heat treatment, results in microstructures that differ from that of the parent alloy. Such microstructures, present in a conventional blade material such as a superalloy, may cause the blade to require excessively frequent repairs in advanced designs that rely on the benefits of directional solidification or single crystal processing to maintain performance.
Materials are characterized by several properties to aid in determining their suitability for use in demanding applications such as gas turbine blades. “Melting temperature” is used herein to refer to the temperature at which liquid metal begins to form as the material is heated. The term “creep life” is used in the art to refer to the length of time until a standard specimen of material extends to a preset length or fractures when subjected to a given stress level at a given temperature. Similarly, the term “fatigue life” is used in the art to describe the length of time until a standard specimen fractures when subjected to a given set of fatigue parameters, including minimum and maximum stress levels, frequency of loading/unloading cycle, and others, at a given temperature. The term “oxidation resistance” is used in the art to refer to the amount of damage sustained by a material when exposed to oxidizing environments, such as, for example, high temperature gases containing oxygen. Oxidation resistance is generally measured as the rate at which the weight of a specimen changes per unit surface area during exposure at a given temperature. In many cases, the weight change is measured to be a net loss in weight, as metal is converted to oxide that later detaches and falls away from the surface. In other cases, a specimen may gain weight if the oxide tends to adhere to the specimen, or if the oxide forms within the specimen, underneath the surface, a condition called “internal oxidation.” A material is said to have “higher” or “greater” oxidation resistance than another if the material's rate of weight change per unit surface area is closer to zero than that of the other material for exposure to the same environment and temperature. Numerically, oxidation resistance can be represented by the time over which an oxidation test was run divided by the absolute value of the weight change per unit area.
Materials particularly noted for high creep life include oxide dispersion strengthened (ODS) materials and directionally solidified eutectic (DSE) alloys. Several materials from these classes have creep lives about three times those measured for conventional superalloys. ODS materials use mechanical techniques during processing to evenly distribute hard oxide particles of sizes less than about 0.1 micron within a metallic matrix, with the particles serving to make deformation of the material more difficult. DSE alloys are characterized by carefully controlled chemistry and processing, which produce a unique microstructure comprising the inherent fibrous or, in some cases, lamellar structure of the eutectic phase, with the fibers or lamellae aligned along a desired axis of the cast part in a manner analogous to a fiber-reinforced composite. DSE materials are also notable for excellent fatigue life, with certain alloys having about three times the fatigue lives measured for conventional superalloys. The careful processing controls needed to produce ODS and DSE alloys cause these materials to be prohibitively expensive.
The so-called “platinum group” of metal elements comprises rhodium (Rh), osmium (Os), platinum (Pt), iridium (Ir), ruthenium (Ru), palladium (Pd), and rhenium (Re)elements noted for high chemical resistance and very high melting temperatures in comparison to conventional superalloys. Several elements from this group are noteworthy as examples of materials with substantially higher oxidation resistance relative to current blade materials. Some platinum group metals and several alloys based on platinum group metals possess excellent resistance to oxidation at temperatures exceeding the capabilities of many Ni-based superalloys. The class of materials referred to as “refractory superalloys” offer additional strength over the platinum group metals, though at the expense of some oxidation resistance. These alloys are based on Ir or Rh, with transition metal additions of up to about 20 atomic percent, and are strengthened by a precipitate phase of generic formula M3X, where M is Rh or Ir and X is typically Ti, V, Ta, or Zr, or combinations thereof. Some alloys of this type can withstand 1-2 hour exposures to at least about 1600° C. without catastrophic oxidation. Creep life and fatigue life data for these alloys are not readily available currently, but the high strength of these alloys suggests they are superior to some degree over conventional superalloys in both creep life and fatigue life at the temperatures and stress levels relevant to gas turbine blade components, although not to the same degree as the best ODS and DSE alloys.
Platinum group metals also have been incorporated into conventional superalloy compositions to produce a class of alloys, herein referred to as “platinum-group metal modified superalloys”, having enhanced oxidation resistance and comparable mechanical properties to conventional superalloys. Typical alloys of this class comprise a conventional superalloy composition to which is added up to about 7 atomic percent of a platinum group metal, such as Ir, Rh, Pt, Pd, and Ru. Use of materials incorporating platinum-group metals has been limited to date due to the high density and very high cost of these materials in comparison to more conventional blade materials.